| Compressor is a key component of aviation gas turbine engine, and in the process of designing compressor designers always obey the rule of circumferentially uniform structure. However, a reasonable aerodynamic explanation about the rule has never been given. The purpose of this paper is to break the inertia of circumferentially uniform structure design and explore the influence of circumferentially non-uniform structure on the compressor’s aerodynamic stability through the analysis of unsteady flow.To reveal the formation and propagation mechanism of rotating stall induced by blade boundary layer separation, and confirm the CFD calculation method of rotating stall, this paper presents a full-annulus numerical simulation in the first stage and the first-stage rotor of a two stages low-speed axial flow compressor. The dynamic process from the stable condition to the rotating stall condition has been simulated. The result shows that the stall inception in the single stage is modal wave, a large scale disturbance, which corresponds to a kind of velocity disturbance produced by circumferentially non-uniform boundary layer separation on static blades. And the rotating frequency of the single full developed stall cell has good coincidence with the experimental result. On the other hand, the stall inception in the isolated rotor blade row is similar to the two-dimensional spike disturbance, belonging to a small scale disturbance. Eventually there exist three stall cells with full development. The rotating frequency of them is higher than the result in the single result and the circumferentially occupied range of them is far less than the single-stage result.To research the influence of non-uniform structure on the performance and aerodynamic stability of compressor, this paper analyses the pitch and stagger non-uniform structure by numerical method. Firstly, the compressor’s characteristic has been obtained through steady numerical simulation. Then, the dynamic stall inception has been obtained through unsteady numerical simulation. According to the calculation results, an effect of local flow redistribution in several passages, which is caused by both kinds of non-uniform structure, becomes even more remarkable near the stall point. The pitch non-uniform structure adversely affects the compressor’s performance and aerodynamic stability, and reduces the compressor’s normal operation range. When the decrease of one blade’s stagger angle is less than 10 degree, the stagger non-uniform structure has a little influence on the compressor’s performance, and the aerodynamic stability is increased. In project 6, an inherent disturbance produced by this stagger non-uniform structure can stop the two-dimensional spike propagating. After the compressor entering rotating stall, the blade whose stagger angle has been reduced can decrease the scope and strength of stall cells, but does not have the capability of suppressing the behavior that stall cells keep rotating circumferentially; when the decrease of one blade’s stagger angle is over 10 degree, the stagger non-uniform structure has negative influence on both the performance and the aerodynamic stability. |
